The present invention relates generally to gas turbine engines, and, more specifically, to turbines therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. Energy is extracted from the gases in turbine stages which power the compressor and a shaft that drives a fan in an aircraft turbofan engine application, or powers a gearbox in a turboshaft application.
A high pressure turbine (HPT) directly follows the combustor and receives the hottest gases therefrom from which energy is initially extracted. A low pressure turbine (LPT) follows the HPT and extracts additional energy from the gases.
Each turbine stage includes a turbine nozzle that preferentially channels the combustion gases to a corresponding row of turbine blades. The nozzle includes hollow stator vanes, and the rotor blades are similarly hollow, for channeling cooling air therethrough during operation.
Each turbine nozzle is an annular assembly of arcuate nozzle segments which must be precisely mounted in the engine coaxially with the axial centerline axis thereof.
In one configuration, the first stage LPT nozzle also includes fairing segments alternating with vane segments. Each fairing segment includes a hollow fairing through which radially extends a structural strut or service lines or conduits for internal engine components.
The nozzle segments may be accurately supported from a surrounding outer casing by corresponding supporting hooks mounted in supporting hangers. Each hook has an axially extending rail which is mounted in an axially extending groove in the support hanger.
This tongue and groove supporting arrangement facilitates assembly of the full row of nozzle segments in the engine. And, the row of nozzle segments is accurately supported and restrained both axially and radially, and coaxially about the axial centerline axis of the engine.
However, during operation the combustion gases flowing through the turbine nozzle heat the nozzle components and cause thermal expansion thereof.
The nozzle, including its airfoil vanes, is suitably cooled during operation by circulating therethrough a portion of pressurized air bled from the compressor.
The cooling air may enter the nozzle and its airfoils from the radially outer band which in turn creates a thermal gradient radially inwardly.
The surrounding support hanger is therefore cooled greater than the outer band, and the supporting hooks thermally expand differently than the thermal expansion of the supporting hangers.
Accordingly, sufficient clearance must be provided in the hanger grooves to accommodate differential thermal expansion of the hooks therein, but that clearance then creates undesirable leakage paths for the cooling air.
Insufficient clearance may cause undesirable binding or interference between the hooks and grooves. This in turn can locally increase loads and stress, and may cause misalignment or mispositioning of the nozzle segments relative to the reference axial centerline axis and adversely affect performance of the combustion gases being channeled through the nozzle to the cooperating row of turbine blades.
These problems increase the complexity of turbine nozzle design, and require suitable solutions therefor.
To reduce undesirable air leakage, auxiliary seals are typically provided at either the forward or aft nozzle hook, or both. And, tubular spoolies may be used to constrain cooling airflow into the individual nozzle airfoils.
Accordingly, it is desired to provide a turbine nozzle having improved mounting features for reducing airflow leakage.